Rotary kinetic fluid motors or pumps – With diversely oriented inlet or additional inlet for...
Reexamination Certificate
2000-06-08
2002-01-22
Look, Edward K. (Department: 3745)
Rotary kinetic fluid motors or pumps
With diversely oriented inlet or additional inlet for...
C415S138000, C415S139000, C415S176000
Reexamination Certificate
active
06340285
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to a gas turbine engine cooling component for end rail cooling, and in particular a turbine engine shroud where each shroud segment provides cooling to both the high pressure and low pressure turbine sections of a gas turbine engine. The present invention further relates to a turbine engine subassembly, and in particular a shroud subassembly that uses a pair of such cooling segments in combination with at least one discourager and primary spline seal.
To increase the efficiency of gas turbine engines, a known approach is to raise the turbine operating temperature. As operating temperatures are increased, the thermal limits of certain engine components can be exceeded, resulting in material failure or, at the very least, reduced service life. In addition, the increased thermal expansion and contraction of these components adversely affects clearances and their interfitting relationships with other components of different thermal coefficients of expansion. Consequently, these components should be cooled to avoid potentially damaging consequences at elevated operating temperatures.
It is common practice then to extract from the main airstream a portion of the compressed air from the compressor for cooling purposes. So as not to unduly compromise the gain in engine operating efficiency achieved through higher operating temperatures, the amount of extracted cooling air should be held to a small percentage of the total main airstream. This requires that the cooling air be utilized with the utmost efficiency in maintaining the temperatures of these components within safe limits.
A particularly important component subjected to extremely high temperatures is the shroud located immediately downstream of the high pressure turbine nozzle, immediately downstream from the combustor. The shroud closely surrounds the rotor of the high pressure turbine and thus defines the outer boundary (flow path) of the extremely high temperature main (hot) gas stream flowing through the high pressure turbine. To prevent material failure and to maintain proper clearance with the rotor blades of the high pressure turbine, adequate shroud cooling is an important concern.
Shroud cooling can be achieved by impingement cooling of the back surface of the shroud, as well as cooling holes that extend from the back surface of the base of the shroud and through to the forward or leading edge of the shroud, the bottom or inner surface of the base in contact with the main (hot) gas stream, and the aft or trailing edge of the shroud to provide both convection cooling inside the holes, as well as impingement and film cooling of the shroud. Cooling flow is also provided through the side panels or rails as convection cooling inside the cooling passages or holes, as well as impingement cooling as cooling air exits from the holes. See, for example, commonly assigned U.S. Pat. No. 5,169,287 (Proctor et al), issued Dec. 8, 1992, which shows a prior embodiment of shroud cooling of the high pressure turbine section of a gas turbine engine. This cooling minimizes local oxidation and burning of the shrouds near the hot main or core gas stream in the high pressure turbine section. Indeed, the cooling holes that exit through the side panel of the shroud of commonly assigned U.S. Pat. No. 5,169,287 can provide important impingement cooling to the side panel of the adjacent shroud.
The leading edge of the shroud is subject to the hottest flow path gas or air, and has the highest heat transfer coefficient, making this section one of the most difficult to cool. As also shown in commonly assigned U.S. Pat. No. 5,169,287, a circumferential row of holes can be angled to also exit at the leading edge of the shroud to provide both convection and film cooling at the leading edge of the shroud. As this cooling film decays and mixes with the hot flow path air, additional circumferential rows of cooling holes can be required to provide more convection and film cooling.
Another type of shroud assembly for a different type of gas turbine engine is shown in commonly assigned U.S. Pat. No. 5,127,793 (Walker et al), issued Jul. 7, 1992. As shown particularly in FIGS. 4 and 4c of U.S. Pat. No. 5,127,793, this prior shroud assembly uses single-piece shroud segments
30
that are designed to span over both the high pressure and low pressure turbine sections of the gas turbine engine. As shown particularly in
FIG. 4
, cooling is provided by directing a portion of the cooling air
74
through ports
78
and through segmented impingement baffles
80
and against the high pressure portion
83
of shroud segment
30
. Another portion of this air
74
is directed into cavity B, with most of it being delivered to cavity C located adjacent the low pressure portions
85
of each shroud segment
30
through holes
84
formed in the support cone portion
86
of turbine shroud support
44
. An impingement baffle
81
attached to shroud support
44
directs and meters impingement cooling air from cavity C onto the low pressure portion
85
of shroud segment
30
. While this prior shroud design of U.S. Pat. No. 5,127,793 provides significant impingement cooling to the back surface of shroud segment
30
in both the high and low pressure sections, it provides no impingement cooling to the side panels or rails of adjacent shroud segments.
The shroud assembly shown in commonly assigned U.S. Pat. No. 5,127,793 extends from approximately the aft end of the upstream turbine nozzle to approximately the leading edge of the downstream turbine nozzle and encloses (i.e., provides a 360° annular structure around) the outer air flow path of a gas turbine engine that typically has a turning nozzle to direct the air flow properly into the blade row, then into a row of blades in the HPT section, and then into another row of blades in the LPT section. Axial gaps between these shroud segments allow for thermal growth over the large range of temperatures the gas turbine engine produces. As hot flow path air passes through the row of turbine blades, work is extracted from the air, thus creating a pressure and temperature drop axially through the blade row. As a result, both the pressure and temperature is higher at the leading edge of the shroud and lower at the trailing edge of the shroud.
A typical sealing method along the axial split lines or gaps between shroud segments is to provide a machined groove or slot in which a thin metal seal (usually referred to as a “spline seal”) is placed, with pressure loading across the seal to provide positive sealing and to minimize air leakage. See FIG 11
a
of commonly assigned U.S. Pat. No. 5,127,793 which shows a pair of longitudinally extending slots in shroud segment
30
, the lower slot receiving the lower or “discourager” spline seal, the upper slot(s) receiving the upper or “primary” spline seal(s). The portion of the axial segment gap that is set up between the shroud segments below the “discourager” seal (commonly referred to as the “trench”) also has hot flow path air traveling axially down it due to the pressure gradient produced by the turbine blade row. Typically no preferential cooling is added to this “trench.” Instead, in the past, air that leaks around the “discourager” seal and the conduction from adjacent metal has been deemed sufficient to cool the axial split lines, i.e., at the side rails or panels of the shroud segments. However, in more recent gas turbine engines that operate at higher temperatures, it has been discovered that oxidation and loss (melting) of the parent material along the axial split-lines of shroud segments can occur.
Accordingly, it would desirable, therefore, to provide a shroud and resulting shroud assembly, particularly for the combined high pressure and low pressure turbine sections, that creates effective impingement cooling for the side panels of adjacent shroud segments. It would also be desirable to provide such impingement cooling while efficiently utilizing the total available cooling air so as not to significantly decrease th
Doughty Roger Lee
Gonyou Craig Alan
Shelton Monty Lee
General Electric Company
Hess Andrew C.
Look Edward K.
Nguyen Ninh
Young Rodney M.
LandOfFree
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